National Advisory Committee for Aeronautics NO. 92 R esearc h Abstracts NOVEMBER 15, 1955 CURRENT NACA REPORTS NACA Kept. 1175 EFFECT OF VARIABLE VISCOSITY AND THERMAL CONDUCTIVITY ON HIGH-SPEED SLIP FLOW BE- TWEEN CONCENTRIC CYLINDERS. T. C. Lin and R. E. Street, University of Washington. 1954. ii, 36p. dia^rs. (NACA Rept. 1175. Formerly TN 2895) The dillerential equations ol slip flow, including the Burnett terms, were first solved by Schamberg as- suming that the coefficients of viscosity and heat con- duction of the gas were constants. The problem is solved herein for variable coefficients of viscosity and thernjal conductivity by applying a transforma- tion leading to an iteration method. The method, starting with the solution for constant coefficients, enables one to approximate the solution for variable coefficients very closely after one or two steps. Satisfactory results are shown to follow from Schaniberg's solution by using his values of the constant coefficients multiplied by a constant factor n, leading to what are denoted as the effec- tive coefficients of viscosity and thermal conduc- tivity. NACA Rept. 1189 THEORETICAL AND EXPERIMENTAL ANALYSIS OF LOW-DRAG SUPERSONIC INLETS HAVING A CIRCULAR CROSS SECTION AND A CENTRAL BODY AT MACH NUMBERS OF 3. 30, 2. 75, AND 2.45. Antonio Ferri and Louis M. Nucci. 1954. ii, 37p. diagrs., photos. (NACA Rept. 1189. Supersedes RM L8H13) Contains theoretical and experimental analysis of circular inlets having a central body at Mach num- bers of 3.30, 2.75, and 2.45. The inlets have been designed in order to have low drag and high pressure recovery. The pressure recoveries obtained are of the same order of magnitude as those previously obtained by inlets having very large external drag. NACA Rept. 1198 A THEORETICAL STUDY OF THE EFFECT OF FORWARD SPEED ON THE FREE-SPACE SOUND- PRESSURE FIELD AROUND PROPELLERS. I. E. Garrick and Charles E. Watkins. 1954. ii, 16p. diagrs., tab. (NACA Rept. 1198. Supersedes TN 3018) The sound-pressure field due to thrust and torque of a propeller in flight at uniform subsonic speed is analyzed by use of a distribution of acoustic doublets located at the propeller disk. The basic element used to synthesize the field is the pressure field of a concentrated force moving uniformly at subsonic speeds, for which an expression generalizing one of Lamb's for the fixed concentrated force is given. The results can be regarded as an extension of the work for the static propeller given by Gutin (NACA TM 1195) for the far pressure field and given by Hubbard and Regier (NACA Rep. 996) for the field near the tips of the rotating propeller. The extendfed formulas are used to calculate the sound field for a, specific two-blade propeller operating at constant \ torque for various forward-speed Maqh,fl\ynbers. NACA Rept. 1209| DEVELOPMENT OF TURBULENCE-MEASURING EQUIPMENT. Leslie S. G. Kovasznay, John Hopkins University. 1954. ii, BOp. diagr§.. photos., tab. (NACA Rept. 12091 Supersedeg,. TN 2839) '"S^ Hot-wire turbulence-measuring equipment has been developed to meet the more-stringent requirements involved in the measurement of fluctuations' in flow parameters at supersonic velocities. The higher mean speed necessitates the resolution of higher fre- quency components than at low speed, and the rela- tively loA' turbulence level present at supersonic speed makes necessary an improved noise level for the equipment. The equipment covers the frequency range from 2 to 70,000 cycles per second. The equipment is adaptable to all-purpose turbulence vvork with improved utility and accuracy over that of older types of equipment. Sample measurements are given to demonstrate the performance. NACA RM E55I16 SPARK IGNITION OF FLOWING GASES. V - AP- PLICATION OF FUEL-AIR-RATIO AND INITL^L- TEMPERATURE DATA TO IGNITION THEORY. Clyde C. Swett, Jr. November 1955. 19p. diagrs. (NACA RM E55I16) Data showing the effect of fuel-air ratio and initial temperature on spark-ignition energy are presented and applied to a previously developed theory of igni- tion. The initial-temperature data are consistent with the theory; fuel-air-ratio data are only par- tially consistent. Probable reasons for the discrep- ancy are discussed. •AVAILABLE ON LOAN ONLY. ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1512 H ST., NW,, THE REPORT TITLE AND AUTHOR. WASHINGTON 25, D C , CITING CODE NUMBER ABOVE EACH TITLE, NACA TM 1388 GENERAL SOLUTIONS OF OPTIMUM PROBLEMS IN NONSTATIONARY FUGHT. (Soluzioni General! di ProWemi dt Ottimo in Volo Non-Stazionario). Angelo Miele. October 1955. 25p. dlagrs., tab. (NACA TM 1388. Trans, from L'Aerotecnica, v.32, no.3, 1952, p. 135-142) A general method concerning optimum problems in nonstationary flight is developed and discussed. Various conditions of flight in a vertical plane (climb with minimum time, climb with minimum fuel consumption, steepest climb, descending and gliding flight with maximum time of space) are studied; the corresponding best techniques of flight, that is, the optimum speed-height relationships, are deter- mined. NACA TN 3415 A UNIVERSAL COLUMN FORMULA FOR LOAD AT WHICH YIELDING STARTS. L. H. Donnell and V. C. Tsien, Illinois Institute of Technology. October 1955. 48p. diagrs., photos., tab. (NACA TN 3415) An analysis is presented of the load at which yielding first occurs in actual columns, taking adequately into account all the factors which have an important ef- fect upon this load. The results are expressed as a formula or chart applicable to all cases. NACA TN 3476 CALCULATED SPANWISE LIFT DISTRIBUTIONS AND AERODYNAMIC INFLUENCE COEFFICIENTS FOR SWEPT WINGS IN SUBSONIC FLOW. Franklin W. Diederich and Martin Zlotnick. October 1955. 173p. diagrs., tabs. (NACA TN 3476) Spanwise lift distributions have been calculated for 61 swept wings with various aspect ratios and taper ratios and with a variety of angle-of-attack or twist distributions, including flap and aileron deflections, by means of the Weissinger method with eight control points on the semispan. Also calculated were aero- dynamic influence coefficients which pertain to a cer- tain definite set of stations along the span. NACA TN 3494 SOUND PROPAGATION INTO THE SHADOW ZONE IN A TEMPERATURE -STRATIFIED ATMOSPHERE ABOVE A PLANE BOUNDARY. David C. Pridmore- Brown and Uno Ingard, Massachusetts Institute of Technology. October 1955. 57p. diagrs., photo. (NACA TN 3494) A theoretical and ejqierimental study of the sound field about a point source over a plane boundary in the presence of a vertical temperature gradient has been made. Methods are presented for analyzing the effects of temperature gradients on the attenuation of sound in the shadow zone of a sound field. NACA RESEARCH ABSTRACTS NO. 92 NACA TN 3495 FAILURE OF MATERIALS UNDER COMBINED RE- PEATED STRESSES WITH SUPERIMPOSED STATIC STRESSES. George Sines, University of California at Los Angeles. November 1955. 69p. diagrs., photos., tabs. (NACA TN 3495) Experiments on biaxial alternating stresses and simple combinations of static stress with alternating stress are reviewed. A general criterion for the effect of static stress on the permissible amplitude of alternating stress is proposed and compared with results of tests performed under more complex stress states. Tests were performed to determine the effect of static compression on alternating tor- sion and the results are compared with the general criterion. A modification of Orowan's theory of fatigue to include the effect of static stress is pre- sented. NACA TN 3531 PILOT'S LOSS OF ORIENTATION IN INVERTED SPINS. Stanley H. Scher. October 1955. lOp. diagrs., photos. (NACA TN 3531) The rising problem of pilot orientation during spins, especially during unintentional inverted spins, is discussed. The free-spinning-tunnel results and reported airplane e.xperiences concerning inverted spins and recoveries are reviewed. Information is provided regarding the nature of inverted spins, optimum control technique for recovery, and some of the factors which apparently contribute to the pilot's loss of orientation. A spin-simulator rig which was recently constructed at the Langley Labo- ratory for use in an attempt to understand better the problems confronting the pilot of a spinning airplane is described. NACA TN 3537 HELICOPTER INSTRUMENT FUGHT AND PRECI- SION MANEUVERS AS AFFECTED BY CHANGES IN DAMPING IN ROLL, PITCH, AND YAW. James B. Whitten, John P. Reeder, and Aimer D. Grim. November 1955. 14p. diagrs., photos. (NACA TN 3537) The damping in roll, pitch, and yaw of a single-rotor helicopter was varied by means of electronic compo- nents, and these variations were evaluated by per- forming instrument approaches and other precision maneuvers. Increased damping In roll was found to be particularly beneficial, whereas corresponding changes In yaw and pitch were less effective. Some operational aspects of helicopter Instrument ap- proaches are also included In the discussion. NACA RESEARCH ABSTRACTS NO. 92 NACA TN 3539 SOME EFFECTS OF SYSTEM NONLINEARITIES IN THE PROBLEM OF AIRCRAFT FLUTTER. Donald S. Woolston, Harry L. Runyan, and Thomas A. Byrdsong. October 1955. 20p. diagrs., tabs. (NACA TN 3539) This paper presents the results of a preliminary in- vestigation of the effect of nonlinear structural terms on the flutter of a two-degree-of-freedom system. The three types of nonlinearities investi- gated were a flat spot, hysteresis, and a cubic spring. Calculations were made on an analog com- puter. For one case, the flat spot, an experimental investigation was also made and good correlation with theory was found. NACA TN 3542 ANALYSIS OF STRESSES IN THE PLASTIC RANGE AROUND A CIRCULAR HOLE IN A PLATE SUB- JECTED TO UNIAXIAL TENSION. Bernard Budiansky and Robert J. Vidensek. October 1955. 39p. diagrs., tabs. (NACA TN 3542) An approximate theoretical solution is presented for the stresses in the plastic range around a circular hole in an infinite sheet subjected to uniaxial tension. The solution is based on the simple deformation theory of plasticity and is found by application of a variational principle in conjunction with the Rayleigh- Ritz procedure and the use of a high-speed computing machine (SEAC). Numerical results are obtained for four different materials, which are characterized by four distinct uniaxial stress-strain curves. The results for stress concentration factor in the plastic range are compared with those obtained from a for- mula due to Stowell. NACA TN 3544 COMPARISON BETWEEN THEORETICAL AND EX- PERIMENTAL STRESSES IN CIRCULAR SEMI- MONOCOQUE CYLINDERS WITH RECTANGULAR CUTOUTS. Harvey G. McComb, Jr., and Emmet F. Low, Jr. October 1955. 20p. diagrs. (NACA TN 3544) Comparisons are made between a theory for calcu- lating stresses about rectangular cutouts in circular cylinders of semimonocoque construction published in NACA TN 3200 and previously published NACA experimental data. The comparisons include stresses in the stringers and shear stresses in the center of the shear panels in the neighborhood of the cutout. The theory takes into account the bending flexibility of the rings in the structure, and this factor is found to be important in the calculation of stresses about cutouts. In general, when the ring flexibility is considered, good agreement is exhibited between the calculated and experimental results. NACA TN 3550 MEASUREMENTS OF THE EFFECT OF TRAILING- EDGE THICKNESS ON THE ZERO- LIFT DRAG OF THIN LOW-ASPECT-RATIO WINGS. John D. Morrow. November 1955. lip. diagrs., photo. (NACA TN 3550. Supersedes KM L50F26) Results of an exploratory free-flight investigation at zero lift of several rocket -powered drag-research models having 4-percent-thlck wings of taper ratio 0.423 are presented for a Mach number range of 0.7 to 1.6. Four wings having trailing edges of different thicknesses were tested. The drag of all the models was measured and is compared with calculated values. NACA TN 3557 A THEORETICAL ANALYSIS OF THE FIELD OF A RANDOM NOISE SOURCE ABOVE AN INFINITE PLANE. Peter A. Franken, Massachusetts Insti- tute of Technology. November 1955. 20p. diagrs. (NACA TN 3557) The sound field about a random noise source above a plane as measured by a receiver with finite band width is studied theoretically. For simplicity, only the far field is considered. The special case of a perfectly reflecting plane is discussed first and the analysis is then extended to include the case of a plane of arbitrary impedance. NACA TN 3567 STUDY OF SCREECHING COMBUSTION IN A 6-INCH SIMULATED AFTERBURNER. Perry L. Blackshear, Warren D. Rayle, and Leonard K. Tower. October 1955. 58p. diagrs., photos., tab. (NACA TN 3567) The mode of oscillation in a screeching 6-inch- diameter simulated afterburner is identified through axial, circumferential, and diametric surveys of sound amplitude and phase. This mode is found to be the first transverse (sloshing) mode in the hot gases downstream of the flame-holder. The devel- opment of a microphone probe suitable for use in screeching combustors is described. This develop- ment includes a theoretical and experimental treat- ment of the attenuation of high-amplitude sound in tubes. NACA TN 3572 AMPLITUDE OF SUPERSONIC DIFFUSER FLOW PULSATIONS. William H. Sterbentz and Joseph Davids. October 1955. 23p. diagrs. (NACA TN 3572. Supersedes RM E52I24) A theoretical method for evaluating the stability characteristics and the amplitude and frequency of pulsation of ram -jet engines without heat addition is presented. Theory and experiment show that the pulsation amplitude of a high-mass-flow-ratio diffuser increases with decreasing mass flow. The theoretical trends for changes in amplitude, frequen- cy, and mean pressure recovery with changes in plenum -chamber volume were experimentally con- firmed. For perforated, convergent-divergent -type diffusers, theory and experiment show the existence of a stability hysteresis loop on the pressure- recovery, mass-flow-ratio curve. NACA TN 3573 EFFECT OF EXHAUST-NOZZLE EJECTORS ON TURBOJET NOISE GENERATION. Warren J. North and Willard D. Coles. October 1955. 26p. diagrs., photo. (NACA TN 3573) Engine noise levels and jet velocity profiles have been obtained with several turbojet exhaust-nozzle ejectors. An insignificant reduction in total sound power was realized. At subsonic nozzle pressure ratios, total sound power from exhaust -nozzle ejectors or bypass exit configurations can be calcu- lated from primary-jet parameters only. NACA TN 3575 BURNING VELOCITIES OF VARIOUS PREMDCED TURBULENT PROPANE FLAMES ON OPEN BURNERS. Paul Wagner. October 1955. 32p. diagrs., photos., tab. (NACA TN 3575) Turbulent burning velocities were measured as a function of Reynolds number for open propane flames. Flames of propane and oxygen diluted with nitrogen, argon, or helium were studied in a variety of burners up to a maximum pipe Reynolds number of 26,000. The ratio of turbulent to laminar burning velocity correlates with the cold-flow Reynolds num- ber for systems of a given diluent. This ratio also correlates with a Reynolds number calculated from values of the turbulent intensity measured at the burner exit. NACA RESEARCH ABSTRACTS NO. 92 BRITISH REPORTS N-40036* Royal Aircraft Establishment (Gt. Brit.) LOW SPEED WIND TUNNEL CALIBRATION OF A Mk. 9A PITOT -STATIC HEAD. J. E. Nethaway. March 1955. 8p. diagrs., tab. (RAE Tech. Note Aero 2364) This note describes the calibration of a Mark 9A pitot-static head in the No. 2 11-1/2 foot wind tunnel. The results show the variation of pitot and static-pressure error coefficients with incidence, at constant tunnel speed. N-40039* Royal Aircraft Establishment (Gt. Brit.) THE DETERMINATION OF THE RELATIVE REAC- TIVITIES OF BIFUNCTIONAL MONOMERS. PART I - THE NATURE OF THE PROBLEM. G. M. Bristow. July 1955. 6p. tab. (RAE Tech. Note Chem. 1256) An account is given of the use of bifunctional mono- mers in the production of cross-linked polymers, and of some methods by which the relative reactivi- ties of the two unsaturated groups can be determined. N-40042* Royal Aircraft Establishment (Gt. Brit.) A LOW-PRESSURE MICRO-ANALYTICAL METHOD FOR DETERMINING OXYGEN, HYDROGEN AND NITROGEN IN METALS. H. C. Davis and J. A. Gray. March 1955. 14p. diagrs., tabs. (RAE Met. 86) A low-pressure microanalytical method for the determination of oxygen, hydrogen, and nitrogen in metals is described. Gases evolved by vacuum- fusion are analyzed physically. The results obtained for oxygen and hydrogen are closely reproducible, but those for nitrogen are less so. NACA RESEARCH ABSTRACTS NO. 92 N-40044* Royal Aircraft Establishment (Gt. Brit.) THE THEORETICAL WAVE DRAG OF OPEN NOSE AXISYMMETRICAL FOREBODIES WITH VARYING FINENESS RATIO, AREA RATIO AND NOSE ANGLE. J. H. Willis and D. G. Randall. February 1955. 33p. diagrs., tab. (RAE Tech. Note Aero 2360) Existing results for the wave drag of open-nose axisymmetrical forebodies are for bodies whose profiles are straight lines or parabolic arcs. These results are here extended to a family of profiles which includes the straight line and the parabolic arc as special cases. Slender body theory is employed throughout. UNPUBLISHED PAPERS N-40045' Royal Aircraft Establishment (Gt. Brit.) TESTS ON HEAT AND CORROSION RESISTING COATINGS FOR MAGNESIUM ALLOYS. J. Mackay and H. G. Cole. June 1955. 9p. tabs. (RAE Tech. Note Chem. 1254) In tests on painted magnesium alloy specimens heated to 200° C at three-monthly intervals during an intermittent seawater spray test, the best com- binations of protective efficiency with hardness and flexibility were given by an epoxy base scheme and by a standard stoving scheme to D. T. D. 235. Other epoxy schemes gave good protection but were relatively inflexible. A standard air-drying scheme toD.T.D. 260A gave very good protection. Butyl titanate paints and a silicone paint gave poor pro- tection. A zinc chrome sealed anodic treatment gave better results as a protective pretreatment than chromate bath iii of D.T.D. 911A. Butyl titan- ate paints broke down on pure aluminum exposed to the same test. N-26685* EXPERIMENTAL DETERMINATION OF THE EF- FECTIVE WIDTH OF FLAT PLATES IN THE ELAS- TIC AND PLASTIC RANGE. (De experimentele bepaling van de meedragende breedte van vlakke platen in het elastische en het plastische gebied). J. F. Besseling. September 1955. lOOp. diagrs., photos., tab. (Trans, from Nationaal Luchtvaart- laboratorium, Amsterdam, S.414, February 1953) Influence of boundary conditions and transverse stiffeners on the effective width of a plate are dis- cussed. The testing apparatus, measuring equip- ment, and test program are reviewed and test re- sults are presented for aluminum alloy 24ST. N-39572 HEAT-RESISTANT SINTER MATERIALS. (Hochwarmfeste Sinterwerkstoffe). F. Benesovsky. (1955). 6p. diagrs., photo., tab. (Trans, from Werkstoffe und Korrosion, Jour., no. 8/9, Aug. - Sept. 1954, p. 288-290) Ceramal materials, produced by powder metallurgy from a titanium carbide base, are suggested for use in turbine and rocket drives. Other ceramal mate- rials suggested for high temperature use are alloys with boride, silicide, and oxide bases. Titanium boride and zirconium boride show high melting points, are generally brittle, and have poor thermal shock properties. It is believed that metal additions may improve these properties. MISCELLANEOUS NACA Rept. 1135 Errata No. 2 on "EQUATIONS, TABLES, AND CHARTS FOR COMPRESSIBLE FLOW. " Ames Research Staff. 1953. NACA Rept. 1175 Errata on "EFFECT OF VARIABLE VISCOSITY AND THERMAL CONDUCTIVITY ON HIGH-SPEED SLIP FLOW BETWEEN CONCENTRIC CYLINDERS. ' T. C. Lin and R. E. Street. 1954. NACA TN 3454 Errata on "EFFECT OF A DISCONTINUITY ON TURBULENT BOUNDARY-LAYER-THICKNESS PARAMETERS WITH APPLICATION TO SHOCK- INDUCED SEPARATION. " Eli Reshotko and Maurice Tucker. May 1955. NACA TN 3483 Errata on "AN ANALYSIS OF ACCELERATION. AIRSPEED, AND GUST-VELOCITY DATA FROM A FOUR-ENGINE TRANSPORT AIRPLANE IN OPERA- TIONS ON AN EASTERN UNITED STATES ROUTE. * Thomas L. Coleman and Mary W. Fetner. September 1955. NACA TN 3514 Errata on "RESPONSE OF HOMOGENEOUS AND TWO-MATERIAL LAMINATED CYLINDERS TO SINUSOIDAL ENVIRONMENTAL TEMPERATURE CHANGE, WITH APPLICATIONS TO HOT-WIRE ANEMOMETRY AND THERMOCOUPLE PYROME- TRY. " Herman H. Lowell and Norman Patton. September 1955. N-37765* SOME ELEVATED TEMPERATURE STRUCTURAL PROBLEMS OF HIGH-SPEED AIRCRAFT. Richard R. Heldenfels. (Presented to SAE Golden Anniversary National Aeronautic Meeting, Los Angeles, California, October 11-15, 1955) 29p. diagrs. This paper contains a discussion of the problen.s of structural design, such as creep, thermal uuckliiig, thermal stresses, and stiffness reduction, as related to the aerodynamic heating of high-speed aircraft. DECLASSIFIED NACA REPORTS THE FOLLOWING REPORTS HAVE BEEN DECLASSIFIED FROM CONFIDENTIAL, 10/14/55 NACA RM A7J02 THE HIGH-SPEED AERODYNAMIC EFFECTS OF MODIFICATIONS TO THE WING AND WING- FUSELAGE INTERSECTION OF AN AIRPLANE MODEL WITH THE WING SWEPT BACK 35°. Lee E. Boddy and Charles P. Morrill, Jr. February 1£ 1948. 34p. diagrs.. photos. (NACA RM A7J02) Wind-tunnel tests at high subsonic Mach numbers were conducted on an airplane model having swept - back wings. Attempts were made to reduce the interference at the plane of symmetry of the swept- back wing and thus increase its divergence Mach number. Also, tests were made with the wing trailing-edge angle decreased, in an effort to elimi- nate the reversal of aileron hinge moment and wing pitching moment suffered by the true-contour wing at high Mach numbers. NACA RESEARCH ABSTRACTS NO. 92 NACA RM A7J03 THE AERODYNAMIC EFFECTS OF ROCKETS AND FUEL TANKS MOUNTED UNDER THE SWEPT- BACK WING OF AN AIRPLANE MODEL. Lee E. Boddy and Charles P. Morrill, Jr. April 23, 1948. 19p. diagrs. (NACA RM A7J03) The effects of externally mounted rockets and fuel tanks on the aerodynamic characteristics of an air- plane model with sweptback wings are presented in this report. Wind-tunnel tests were made at high subsonic Mach numbers to determine the effect of the external equipment on the drag, the longitudinal stability and control, and the lateral control of the model. NACA RM A50J26 EXPERIMENTAL DAMPING IN PITCH OF 45° TRI- ANGULAR WINGS. Murray Tobak, David E. Reese, Jr., and Benjamin H. Beam. December 1, 1950. 63p. diagrs., photo. (NACA RM A50J26) Results are presented of a wind-tunnel investigation of the variation of the damping-ln-pitch parameter •^m + C_ with Mach number and axis of rotation position for two triangular wings having leading edges swept back 45°, with and without a body. Tests were conducted at subsonic speeds over a Mach number range of 0.23 to 0.94 and at supersonic speeds from 1.15 to 1.70. The measured damping coefficients are compared with theoretical results at both subsonic and supersonic speeds. NACA RM A53J02 PRELIMINARY RESULTS OF AN INVESTIGATION OF THE EFFECTS OF SPINNER SHAPE ON THE CHARACTERISTICS OF AN NACA D-TYPE COWL BEHIND A THREE-BLADE PROPELLER, IN- CLUDING THE CHARACTERISTICS OF THE PRO- PELLER AT NEGATIVE THRUST. Robert M. Reynolds. November 1953. 15p. diagrs., photo., tab. (NACA RMA53J02) Preliminary results of measurements of the ram- recovery ratio at the inlet of an NACA D-type cowl behind an operating propeller in combination with a 1 -series and a modified-conical spinner, maximum efficiency of the propeller with the 1 -series spinner and the spinner-cowling combinations, and the nega- tive thrust characteristics of the propeller at low speeds are summarized. Tests were conducted at Mach numbers from 0.2 to 0.8, for propeller blade angles from 33° to 63°, and for various inlet- velocity and advance ratios. Negative-thrust char- acteristics of the propeller were measured at a Mach number of 0.15 for blade angles from 25° to -20°. All tests were made with the model at an angle of attack of 0° and at a Reynolds number of 1.0 million per foot. NACA RESEARCH ABSTRACTS NO. 92 NACA RM A53J07 THE EFFECTS OF HORIZONTAL-TAIL HEIGHT AND A PARTIAL-SPAN LEADING-EDGE EXTEN- SION ON THE STATIC LONGITUDINAL STABILITY OF A WING-FUSELAGE-TAIL COMBINATION HAVING A SWEPTBACK WING. Angelo Bandettini and Ralph Selan. March 1954. 54p. diagrs., photos., 2 tabs. (NACA RM A53J07) Test results are presented to show the effects of horizontal-tail height (22- and 8-percent semispan above wing chord plane extended) on the static longi- tudinal stability of a model having a wing with 35° sweepback, an aspect ratio 4.5, and a taper ratio 0.5. The model was also modified by a wing-leading-edge chord extension and tested with the tail in the low position. Tests were conducted at various Mach numbers up to 0.92 at a Reynolds number of 2,000,000 and at a Mach number of 0.20 at a Reynolds number of 11,000,000. The results of airstream surveys in the region of the tail are also presented. NACA RM E50J24 DYNAMIC INVESTIGATION OF TURBINE - PROPELLER ENGINE UNDER ALTITUDE CONDI- TIONS. Richard P. Krebs, Seymour C. Himmel, Darnold Blivas, and Harold Shames. December 6, 1950. 55p diagrs., photo. (NACA RM E50J24) An altitude-wind-tunnel investigation of the dynamics of a turbine -propeller engine employing the frequency-response technique was conducted over a range of pressure altitudes from 10,000 to 30,000 feet. The dynamic responses generalized for pres- sure altitudes over the range of frequencies investi- gated. The generalized time constants were found to be approximately 1.0 second for the engine-propeller combination, 0.36 second for the propeller alone, and 2.4 seconds for the engine alone. These values were in good agreement with those predicted from steady - state-performance data. NACA RM E51J11 IGNITION-DELAY CHARACTERISTICS IN MODIFIED OPEN-CUP APPARATUS OF SEVERAL FUELS WITH NITRIC ACID OXIDANTS WITHIN TEMPERA- TURE RANGE 70° TO -105° F. Riley O. Miller. December 1951. 30p. diagrs., 4 tabs. (NACA RM E51J11) Fluid properties and low-temperature ignition delays were obtained for approximately 90 fuel-oxidant com- binations. A red fuming nitric acid containing ap- proximately 3 percent water and 19 percent nitrogen tetroxide froze at approximately -87° F and ignited several low-viscosity fuel blends of aromatic amines in triethylamine at -76° F and lower. With this acid, the following average ignition delays were obtained with a blend of 30 percent o-toluidine in triethyla- mine: Temperature, OF Delay, milliseconds 70 19 -40 24 -76 38 -87 61 -105 210 NACA RM E52J27 PRELIMINARY INVESTIGATION OF A PERFO- RATED AXIALLY SYMMETRIC NOZZLE FOR VARYING NOZZLE PRESSURE RATIOS. Eli Reshotko. January 1953. 43p. diagrs., photo., 2 tabs. (NACA RM E52J27) The performance characteristics of a perforated axially symmetric convergent -divergent nozzle were investigated in an attempt to achieve improved convergent-divergent nozzle thrust performance at below design pressure ratios. The purpose of the perforations was to allow inflow of air into the over- expanded portion of the nozzle, thus advancing sepa- ration of the flow. The flow through the perfora- tions was found to advance separation only when the perforations were liberally placed over the entire divergent portion of the nozzle. A local concentra- tion of perforations caused separation only in the local region of perforation. The use of low energy atmospheric bleed air reduced thrust losses by as much as 50 percent at appreciably overe.xpanded operation. For underexpanded flow, air flowing out through the perforations caused significant thrust loss. With shrouding to prevent this outbleed, thrusts 5 to 10 percent less than those of the un- perforated nozzle were obtained. The use of high energy bleed was unsatisfactory since the inlet momentum penalty of the bleed air was in many cases greater than the additional thrust obtained. NACA RM L8I29 PRELIMINARY RESULTS OF NACA TRANSONIC FLIGHTS OF THE XS-1 AIRPLANE WITH 10- PERCENT-THICK WING AND 8-PERCENT-THICK HORIZONTAL TAIL. Hubert M. Drake, Harold R. Goodman, and Herbert H. Hoover. October 13, 1948. 18p. diagrs., photos. (NACA RM L8I29) Contains results of exploratory flights at altitudes of about 40,000 feet to a maximum Mach number of 1.06. Data are presented showing the longitudinal trim changes, elevator effectiveness in producing acceleration, and rudder effectiveness as a Junction of Mach number. Data on lateral oscillations are also presented. NACA RM L8J12 HIGH-SPEED WIND-TUNNEL INVESTIGATION OF A SWEPTBACK WING WITH AN ADDED TRIANGU- LAR AREA AT THE CENTER. Beverly Z. Henry, Jr. January 14, 1949. 24p. diagrs., tabs. (NACA RM L8J12) Results are presented of an investigation in the Langley 8-foot high-speed tunnel of two sweptback wings of different plan form. The purpose of the investigation was to determine the effects of the addition of a triangular area to the inboard section of a conventional sweptback wing in order to produce a wing employing two stages of sweepback. Lift, drag, and pitching-moment characteristics are pre- sented to Illustrate these effects for a Mach number range of 0.40 through 0.935. NACA RM L8K23 FREE-FLIGHT INVESTIGATION AT TRANSONIC AND SUPERSONIC SPEEDS OF THE ROLLING EF- FECTIVENESS OF SEVERAL AILERON CONFIGU- RATIONS ON A TAPERED WING HAVING 42.7° SWEEPBACK. Carl A. Sandahl. January 11, 1949. 23p. diagrs., photos., tab. (NACA RM L8K23) An investigation was made of several aileron modi- fications in conjunction with a tapered, sweptback wing having circular-arc airfoil sections of rela- tively large thickness ratio. The modified ailerons eliminated the reversal of rolling effectiveness at transonic speeds obtained with the true-contour ai- lerons at small deflections. NACA RM L9F07 PRELIMINARY THEORETICAL AND FLIGHT IN- VESTIGATION OF THE LATERAL OSCILLATION OF THE X-1 AIRPLANE. Hubert M. Drake and Helen L. Wall. July 19, 1949. 24p. diagrs., photo., tab. (NACA RM L9F07) A small -amplitude, undamped, lateral oscillation has been encountered in flight tests of the X-1 airplane. The oscillation occurs in subsonic and supersonic flight, in maneuvers, and power on and off. The calculations indicate that a change, in the positive direction, of the inclination of the principal axis with respect to the flight path should have a considerable stabilizing effect. NACA RM L9G19a MEASUREMENTS OF AILERON EFFECTIVENESS OF THE BELL X-1 AIRPLANE AT MACH NUMBERS BETWEEN 0.9 AND 1.06. Hubert M. Drake. August 4, 1949. 5p. diagrs. (NACA RM L9G19a) Presents results of flight measurements of aileron effectiveness of the X-1 airplane up to a Mach num- ber of 0.94. The data indicate a 75 percent loss of aileron effectiveness between M = 0.82 and M = 0.94. NACA RM L50G20 ELEVATOR-STABILIZER EFFECTIVENESS AND TRIM OF THE X-1 AIRPLANE TO A MACH NUM- BER OF 1.06. Hubert M. Drake and John R. Garden. November 1, 1950. 12p. diagrs. (NACA RM L50G20) The relative elevator -stabilizer effectiveness of the X-1 has been determined to decrease from a value of 0.25 at a Mach number of 0.78 to a value of 0.05 at a Mach number of 1.0. At supersonic speeds the ef- fectiveness increases. The variation between the trim curves at various stabilizer settings is caused by the variation in effectiveness and the fact that the effectiveness is nonlinear at Mach numbers between 0.94 and 0.97. It was found that, with the elevator fixed at zero, only about 0.5° of stabilizer movement would be required to trim through the Mach number range from 0.78 to 1.02. NACA RESEARCH ABSTRACTS NO. 92 NACA RM L50J25 A TRANSONIC -WING INVESTIGATION IN THE LANGLEY 8-FOOT HIGH-SPEED TUNNEL AT HIGH SUBSONIC MACH NUMBERS AND AT A MACH NUM- BER OF 1.2. WING-FUSELAGE CONFIGURATION HAVING A WING OF 60° SWEEPBACK, ASPECT RATIO 4, TAPER RATIO 0.6, AND NACA 65A006 AIRFOIL SECTION. Raymond B. Wood and Frank F. Fleming. January 24, 1951. 43p. diagrs., photo. (NACA RM L50J25) An investigation was conducted in the Langley 8-foot high-speed tunnel of the aerodynamic characteristics of a wing swept back 60° at the quarter chord, with aspect ratio 4, taper ratio 0.6, and an NACA 65A006 airfoil section. The tests were conducted through a Mach number range from 0.6 to 0.96 and at a Mach number of 1.2. Data are presented for a wing fuselage and for a wing with wing-fuselage interfer- ence. Wake-survey-study results and the measure- ments of the angle of downwash for a probable tail location, approximately 38 percent of the wing semi- span above the wing-chord plane, are included. NACA RM L51D17 AN INVESTIGATION OF FOUR WINGS OF SQUARE PLAN FORM AT A MACH NUMBER OF 6.86 IN THE LANGLEY 11-INCH HYPERSONIC TUNNEL. Charles H. McLellan, Mitchel H. Bertram, and John A. Moore. June 1951. 47p. diagrs., photos. (NACA RM L51D17) The results of tests of four wings at a Mach number of 6.86 in the Langley 11-inch hypersonic tunnel are presented. The wings tested had 4-inch square plan forms with 5-percent-thick diamond, half-diamond, wedge, and half-circular-arc sections. The bound- ary layer has been found to have a large effect on the wing pressure distributions. Reasonable agreement was indicated between the aerodynamic coefficients from experimental pressure data and inviscid theory. Total drag measurements showed good agreement with the theory at low angles of attack when the ef- fects of surface friction were included. At the higher angles of attack, both lift coefficient and drag coefficient were found to be slightly below the values predicted by the two-dimensional theory. NACA RM L51J10 INVESTIGATION BY THE TRANSONIC -BUMP METHOD OF A 35° SWEPTBACK SEMISPAN MODEL EQUIPPED WITH A FLAP OPERATED BY A SERIES OF SERVOVANES LOCATED AHEAD OF AND GEARED TO THE FLAP. William H. Phillips and Robert F. Thompson. December 1951. 39p. diagrs., photo. (NACA RM L51J10) Lift. drag, pitching-moment. rolling-moment, and yawing-moment data in the Mach number range from 0.6 to 1.0 obtained from wind-tunnel tests of a low- aspect-ratio sweptback airfoil model with a servo- vane control are presented. The control utilizes the drag force and spoiler action of a set of vanes to deflect a flap-type control. Comparison of lift in- crement and center-of-pressure location is made with previously published data from tests of a con- ventional flap-type control. NACA RESEARCH ABSTRACTS NO. 92 UNIVERSITY OF FLORIDA 1262 08153 288 8 NACA RM L51J30 SUMMARY OF FLUTTER EXPERIENCES AS A GUIDE TO THE PRELIMINARY DESIGN OF LIFTING SURFACES ON MISSILES. Dennis J. Martin. November 1951. 16p. diagrs. (NACA RM L51J30) This report presents a limited review of some ex- periences in flight testing of missiles and of wing flutter investigations that may be of interest in mis- sile design. Several types of flutter which may be of concern in missile studies are briefly described. Crude criteria are presented for two of the most common types of flutter to permit a rapid estimate to be made of the probability of the occurrence of flutter. Many of the details of the flutter picture have been omitted, and only the broader elements have been retained so as to give the designer an overall view of the subject. NACA RM L52D01 A STUDY OF THE FLOW OVER A 45° SWEPTBACK WING-FUSELAGE COMBINATION AT TRANSONIC MACK NUMBERS. Richard T. Whitcomb and Thomas C. Kelly. June 1952. 60p. diagrs., photos. (NACA RM L52D01) Pressure distributions, tuft patterns, and schlieren surveys have been obtained for a 45° sweptback wing-fuselage combination in the Langley 8-foot transonic tunnel at transonic Mach numbers to 1.11 and angles of attack to 20°. The results provide an indication at transonic Mach numbers of the nature of the formation of shock waves on the wing and fuselage, wing-fuselage interference, and the de- velopment of separation and the separation vortex. NACA RM L52J21a INVESTIGATIONS AT SUPERSONIC SPEEDS OF THE BASE PRESSURE ON BODIES OF REVOLUTION WITH AND WITHOUT SWEPTBACK STABILIZING FINS. Eugene S. Love and Robert M. O'Donnell. December 1952. 66p. diagrs., photos. (NACA RM L52J21a) Results are presented from an investigation at Mach numbers of 1.62, 1.93, and 2.41 of the variation with Reynolds number of the base pressure on bodies of revolution at zero lift, with and without sweptback stabilizing fins. Included are the effects of varying nose and base shapes and cutoff length, the effects of the presence of sting supports of varying diam- eter, and the effects of disturbances entering the wake. The overall Reynolds number range was ap- proximately from 1 X 10^ to 10 x 10^. NACA RM L52J23a INVESTIGATION OF THE AERODYNAMIC CHARAC- TERISTICS OF THE NACA RM-10 MISSILE (WITH FINS) AT A MACH NUMBER OF 1.62 IN THE LANGLEY 9-INCH SUPERSONIC TUNNEL. Donald E. Coletti. December 1952. 21p. diagrs. (NACA RM L52J23a) An investigation was made of a 0.050-scale model of the RM-10 missile at a Mach number of 1.62 and a Reynolds number of 2.66 x 10^. Measurements were made of lift, drag, and pitching moment over an angle-of-attack range of t5°. The effects of the ratio of sting-shield diameter to base diameter were also investigated. Comparisons are made with re- sults of tests in other facilities at widely different Reynolds numbers. NACA RM L52K06 PRESSURE DISTRIBUTION AND PRESSURE DRAG FOR A HEMISPHERICAL NOSE AT MACH NUM- BERS 2.05. 2.54, AND 3.04. Leo T. Chauvin. December 1952. 14p. diagrs., photos. (NACA RM L52K06) An experimental investigation of the pressure dis- tributions on a hemispherical nose 3.98 inches in diameter, mounted on a cylindrical support, has been made at Mach numbers of 2.05, 2.54, and 3.04 and for Reynolds numbers of 4.44 x 10^, 4.57 x 10^, and 4.16 x 10^, respectively. The Reynolds number was based on body diameter and free-stream condi- tions. Pressure-drag coefficients were calculated and good agreement was obtained between these tests and other investigations. NACA RM L53E04 AILERON AND ELEVATOR HINGE MOMENTS OF THE BELL X-1 AIRPLANE MEASURED IN TRAN- SONIC FLIGHT. Hubert M. Drake and John B. McKay. June 1953. 27p. diagrs. (NACA RM L53E04) Hinge moments have been measured on the aileron and elevator of the Bell X-1 airplane having the 10- percent-thick wing and 8-percent-thick tail. The aileron measurements were made by means of strain gages and pressure distributions while the elevator measurements were made by means of the wheel-force strain gages. The elevator hinge- moment characteristics were determined to a Mach number of 1. 18 and the aileron hinge moments to a Mach number of 1. 13. NACA RM L53F08 FLIGHT MEASUREMENTS OF LIFT AND DRAG FOR THE BELL X-1 RESEARCH AIRPLANE HAV- ING A 10-PERCENT-THICK WING. Edwin J. Saltzman. September 1953. 37p. diagrs., tab. (NACA RM L53F08) Lift and drag results have been obtained from power- off flight tests of the Bell X-1 (lO-percent-thick wing) airplane for Mach numbers 0.68 to 1.01. Com- parisons of drag are made with 8-percent-thick-wing flight tests and 10-percent-thick-wing wind-tunnel results. 10 NACA RM L53I09a LOW-SPEED WIND-TUNNEL INVESTIGATION OF A JET CONTROL ON A 35° SWEPT WING. John G. Lowry and Thomas R. Turner. October 1953. 9p. diagrs. (NACA RM L53I09a) A low-speed wind-tunnel investigation was made of a jet control that obtains its effectiveness from both the jet reaction and from the change in circulation around the wing due to the jet's acting as a spoiler. The jet control was investigated as an aileron on a 35° sweptback wing of aspect ratio 4.76. The in- vestigation was of exploratory nature and was limited to the case where the jet was supplied with air at stagnation pressure. The results indicated that such a jet could be used as an emergency control. NACA RM L53J01a WIND-TUNNEL INVESTIGATION OF THE EFFECTS OF STEADY ROLLING ON THE AERODYNAMIC LOADING CHARACTERISTICS OF A 45° SWEPT- BACK WING AT HIGH SUBSONIC SPEEDS. James W. Wiggins and Richard E. Kuhn. November 1953. 22p. diagrs., photos. (NACA RM L53J01a) The aerodynamic loading characteristics of a 45 sweptback wing of aspect ratio 4 in combination with a fuselage during steady roll are presented. The tests covered Mach numbers of 0.70, 0.85, and 0.91, and angles of attack up to 13°. The effects of fences at a Mach number of 0.85 and a comparison of measured and calculated load distribution are in- cluded. NACA RM L53J09a WIND-TUNNEL INVESTIGATION AT LOW SPEED OF THE EFFECT OF VARYING THE RATIO OF BODY DIAMETER TO WING SPAN FROM 0.1 TO 0.8 ON THE AERODYNAMIC CHARACTERISTICS IN PITCH OF A 45° SWEPTBACK-WING— BODY COMBINATION. Harold S. Johnson. November 1953. 32p. diagrs., photo., tab. (NACA RM L53J09a) Low-speed lift, drag, and pitching-moment data were obtained for a family of bodies and wing-body combinations to determine the effect of varying the ratio of body diameter to wing span from 0.1 to 0.8. The bodies had ogival noses and cylindrical after- bodies. The untapered 6-percent-thick wings had aspect ratios of 3 and 45° of sweepback. The lift data of the bodies alone and the wing-body combina- tions are compared with several existing theories. NACA RM L53J19 AN EXPERIMENTAL AND THEORETICAL INVES- TIGATION AT HIGH SUBSONIC SPEEDS OF THE EFFECTS OF HORIZONTAL-TAIL HEIGHT ON THE AERODYNAMIC CHARACTERISTICS IN SIDESLIP OF AN UNSWEPT, UNTAPERED TAIL ASSEMBLY. Harleth G. Wiley and Donald R. Riley. December 1953. 71p. diagrs., tab. (NACA RM L53J19) This paper presents the effects at high subsonic speeds of horizontal-tail height on the aerodynamic characteristics in sideslip at 0° angle of attack of an NACA RESEARCH ABSTRACTS NO. 92 unswept, untapered empennage. Configurations in- vestigated included the fuselage alone, fuselage plus vertical tail, fuselage plus horizontal tail, and the fuselage plus vertical tail with the horizontal tail located at 0, 26, 59, and 100 percent vertical- surface span. Tests were made at 0° angle of attack through a sideslip range of -2° to 20° over a Mach number range of 0.50 to 0.94. NACA RM L53J29 WIND-TUNNEL INVESTIGATION AT HIGH AND LOW SUBSONIC MACH NUMBERS OF TWO UNSWEPT WINGS HAVING NACA 2-006 AND NACA 65A006 AIRFOIL SECTIONS. Stanley F. Racisz. December 1953. 40p. diagrs., photo., tab. (NACA RM L53J29) An investigation has been made of two unswept wings with aspect ratios of 4 and taper ratios of 0.2. One wing had airfoil sections designed for high maximum lift at low speeds (NACA 2-006), and the other wing had NACA 65A006 airfoil sections. Each wing was mounted on a slender body of revolution. The lift, drag, and pitching-moment characteristics were determined at Reynolds numbers from 1 x 10^ to 7.5 X 10^ for Mach numbers below 0.2 for the wings with and without split flaps and for the wings with and without leading-edge roughness. The character- istics of the plain wings were also determined for several values of Reynolds number at Mach numbers up to about 0.92. Gains obtainable by the use of the NACA 2-006 airfoil section are evident for Mach numbers up to 0.65 from the comparisons of the re- sults for the two wings. NACA RM L53L08a EXPERIMENTAL CONVECTIVE HEAT TRANSFER TO A 4-INCH AND 6-INCH HEMISPHERE AT MACH NUMBERS FROM 1.62 TO 3.04. Leo T. Chauvin and Joseph P. Maloney. February 1954. ISp. diagrs., photos. (NACA RM L53L08a) Equilibrium temperatures and heat-transfer coeffi- cients for a hemispherical nose have been measured for Mach numbers from 1.62 to 3.04. Heat transfer to the surface of the hemisphere was presented as Stanton number against Reynolds number for various surface heating conditions. Heat transfer at the stagnation point has been measured and correlated with theory. Transition from a laminar to a turbu- lent boundary layer was obtained at Reynolds num- bers of approximately 1 x 10^ corresponding to a region on the body located between 45° and 60° from the stagnation point. NACA RM L53L15 INVESTIGATION OF A PULSE-JET-POWERED HELICOPTER ROTOR ON THE LANGLEY HELI- COPTER TEST TOWER. Edward J. Radin and Paul J. Carpenter. February 1954. 23p. diagrs., photos. (NACA RM L53L15) A helicopter rotor powered by tip-located pulse-jet engines has been investigated on the Langley heli- copter test tower to determine its basic hovering characteristics as well as the power-off drag and propulsive characteristics of the engines. The noise intensity in the vicinity of the pulse-jet engines was also determined. NACA - Langley Field, Va.